Cell structure thermal barrier coating

ABSTRACT

A combustor component of a gas turbine engine includes a thermal barrier coating on a substrate, the thermal barrier coating defines a multiple of cells.

BACKGROUND

The present disclosure relates to a thermal barrier coating, and moreparticularly to a combustor with a thermal barrier coating.

A gas turbine engine includes a compressor for compressing air which ismixed with a fuel and channeled to a combustor wherein the mixture isignited within a combustion chamber to generate hot combustion coregases. At least some combustors include combustor liners to channel thecombustion gases to a turbine which extracts energy from the combustioncore gases to power the compressor, as well as produce useful work topropel an aircraft in flight or to power a load, such as an electricalgenerator.

The combustor liners often include a thermal barrier coating to increasedurability. The difference in properties between the ceramic thermalbarriers and the metal substrates to which the ceramic is applied maylead to mismatched strains which ultimately lead to areas of coatingspallation which may tend to spall or flake. Once spalled, substratedegradation in the form of cracking and oxidation may follow.

SUMMARY

A component of a gas turbine engine according to an exemplary aspect ofthe present disclosure includes a thermal barrier coating on asubstrate, the thermal barrier coating defines a multiple of cells.

A combustor component of a gas turbine engine according to an exemplaryaspect of the present disclosure includes a substrate which defines amultiple of offsets and a thermal barrier coating on the substrate, thethermal barrier coating defines a multiple of cells, each of themultiple of cells correspond with at least one of the multiple ofoffsets.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of a gas turbine engine;

FIG. 2 is a perspective partial sectional view of an exemplary annularcombustor that may be used with the gas turbine engine shown in FIG. 1;

FIG. 3 is a cross-sectional view of an exemplary combustor that may beused with the gas turbine engine shown in FIG. 2;

FIG. 4 is a facial view of a combustor component;

FIG. 5 is a cross-sectional view of one non-limiting embodiment of thecombustor component; and

FIG. 6 is a cross-sectional view of another non-limiting embodiment ofthe combustor component.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath whilethe compressor section 24 drives air along a core flowpath forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. The inner shaft 40 and the outer shaft50 are concentric and rotate about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 54, 46 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

With reference to FIG. 2, the combustor 56 generally includes an outerliner 60 and an inner liner 62. It should be understood that variouscombustor arrangements such as a can combustor as well as other hightemperature components such as turbine components may alternativelybenefit herefrom.

With reference to FIG. 3, outer liner 60 and inner liner 62 are spacedinward from a combustor case 64 such that a combustion chamber 68 isdefined between liners 60, 62. The outer liner 60 and combustor case 64define an outer passageway 70. The inner liner 62 and combustor case 64define an inner passageway 72. Combustion chamber 68 is generallyannular in shape and is defined between liners 60, 62. Outer and innerliners 60, 62 extend toward the turbine section 28.

The outer and inner liners 60, 62 support a multiple of liner panels 74.Each liner panel 74 generally includes a metallic substrate 76 with athermal barrier coating 78 (FIG. 4) on an inner surface 80 which facesthe combustion chamber 68 (FIG. 5). It should be understood thatalthough a particular combustor is illustrated, other combustor typeswith various combustor liner panel arrangements will also benefitherefrom. It should be further understood that the disclosed liner panelis but a single illustrated embodiment and should not be limited onlythereto such that the disclosed liner panel 74 may be considered but onecombustor component of various types manufactured of a substrate 76 uponwhich the thermal barrier coating 78 is applied as disclosed herein.

The substrate 76 may be a nickel base superalloy, other metallicmaterial, or Ceramic Matrix Composite material. The thermal barriercoating 78 may be applied in, for example, a plasma spray coatingprocess in which powders are injected into a high temperature, highvelocity stream of ionized gases. At the point where the powders areinjected into the gas stream, the temperature can be about15,000.degrees F (8315 C). As a result, the powders are typically moltenwhen they strike the surface of the substrate forming an interlocking“splat” type structure. It should be understood that the thermal barriercoating 78 may be sprayed on a bondcoat which has been applied to thesubstrate 76 which has been found to improve adhesion and is well knownin the industry.

With reference to FIG. 4, the thermal barrier coating 78 is applied tothe substrate 76 (FIG. 5) in a manner to form a grid pattern whichincludes a multiple of cells 82 each separated by a narrow gap 84 (alsoillustrate in FIG. 5). The gap 84 may be formed to be exceedingly narrowyet still facilities thermal barrier coating 78 strain tolerance as thegaps 84 define the maximum size of potential ‘mudflat’ cracks that mayoccur due to sintering. This segregation facilitates durability andaccommodation of thermal gradients.

Each cell 82 may be of a particular shape such as hexagonal (shown),square, triangular or other shape. In one non-limiting embodiment, eachcell 82 may be of approximately 0.25 inches (6.35 mm) across opposedcorners of the illustrated hexagonal shape.

The multiple of cells 82 with the respective narrow gap 84 may bemanufactured, for example, with a matrix grid of a polyester fugitivematerial which is applied over the substrate 76 prior to the “splat”type plasma spray process then thereafter baked out to form the gaps 84.The gap 84 could be as narrow as a crack and still provide a measure ofstrain relief and durability improvement such that an alternativethermal process may include a laser testament to pre-treat the ceramicand produce the matrix grid of gaps 84. It should be understood thatvarious processes may alternatively or additionally be utilized toessentially mask or mark the substrate 76 to form the matrix grid ofgaps 84.

An offset 86 interfaces with each of the multiple of cells 82. Theoffset 86 may extend outward from the surface 80 of the substrate 76 toform a post which extends into the thermal barrier coating 78 (FIG. 5).Alternatively, an offset 86′ may extend into the surface 64 of thesubstrate 76′ to form a divot which at least partially receives thethermal barrier coating 78 (FIG. 6). In one disclosed non-limitingembodiment, the offset 86, 86′ defines a reverse taper such as afrustro-conical structure relative to the surface 80 of the substrate 76that is two-thirds (⅔) the thickness of the thermal barrier coating 78to form an interlock for the thermal barrier coating 78 at each cell 82.

The offset 86, 86′ provides a reference point about which sinteringshrinkage of the thermal barrier coating 78 will interlock to facilitateadhesion of the thermal barrier coating 78 to the substrate 76. That is,the thermal barrier coating 78, when sintered, mechanically interlocksonto the post or into the divot. This mechanical interlock significantlyincreases the life of the thermal barrier coating 78 and thereforeincreases the life of the components onto which the thermal barriercoating 78 is applied. As each cell 82 is provided with an offset 86,86′, adhesion to the substrate 76 is supplemented and the thermalbarrier coating 78 may last for the entire life of the component whichis protected thereby.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

1. A component of a gas turbine engine comprising: a substrate; and athermal barrier coating on said substrate, said thermal barrier coatingdefines a multiple of cells.
 2. The component as recited in claim 1,wherein said substrate defines an offset within each of said multiple ofcells.
 3. The component as recited in claim 2, wherein said offsetwithin each of said multiple of cells is a post.
 4. The component asrecited in claim 3, wherein said post within each of said multiple ofcells is frustro-conical.
 5. The component as recited in claim 3,wherein said post is located within a center of each of said multiple ofcells.
 6. The component as recited in claim 3, wherein said post definesa height that is approximately two-thirds the thickness of said thermalbarrier coating.
 7. The component as recited in claim 2, wherein saidoffset within each of said multiple of cells is a divot.
 8. Thecomponent as recited in claim 7, wherein said divot is located within acenter of each of said multiple of cells.
 9. The combustor component asrecited in claim 1, wherein each of said multiple of cells are hexagonalin shape.
 10. The combustor component as recited in claim 1, whereineach of said multiple of cells are separated by a gap.
 11. The combustorcomponent as recited in claim 1, said substrate is a nickel basesuperalloy.
 12. A combustor component of a gas turbine enginecomprising: a substrate which defines a multiple of offsets; and athermal barrier coating on said substrate, said thermal barrier coatingdefines a multiple of cells, each of said multiple of cells correspondwith at least one of said multiple of offsets.
 13. The combustorcomponent as recited in claim 12, wherein at least one of said multipleof offsets is a post.
 14. The combustor component as recited in claim13, wherein said post is frustro-conical.
 15. The combustor component asrecited in claim 13, wherein each of said multiple of posts is locatedwithin a center of each of said multiple of cells.
 16. The combustorcomponent as recited in claim 13, wherein said post is approximatelytwo-thirds the thickness of said thermal barrier coating.
 17. Thecombustor component as recited in claim 12, wherein at least one of saidmultiple of offsets is a divot.
 18. The combustor component as recitedin claim 17, wherein s each of said multiple of divots is located withina center of each of said multiple of cells
 19. The combustor componentas recited in claim 12, wherein each of said multiple of cells arehexagonal in shape.
 20. The combustor component as recited in claim 19,wherein each of said multiple of cells are separated by a gap.